A compressor (or turbine) of a gas turbine engine comprises rotor components, including rotor blades and a rotor drum, and stator components, including stator vanes and a stator casing. The compressor is arranged with a number of alternating rotor blade and stator vane stages as is well known. The efficiency of the compressor is influenced by the running clearances or radial gap between its rotor and stator components. The radial gap or clearance between the rotor blades and stator casing and between the stator vanes and the rotor drum are set to be as small as possible to minimise over tip leakage of working gases.
Typically, the minimum clearance that can be set is determined by the transient variations of the clearance during engine operation. However, when engine demand changes the transient conditions cause the components to experience different thermal gradients and thermal lag. Along with changes of rotational speeds that influence radial position of components and different component materials the clearances are significantly affected during these transient engine conditions. In some locations the rotor and stator components are allowed to lightly touch, transiently, leaving a rub mark that can be seen.
In addition, rotor components can rotate eccentrically or ‘wobble’ about the engine's rotational axis. This eccentricity can contribute to rotor and stator rubbing. In particular, a rotor-stator rub can occur in only a discrete circumferential region. Also, when the engine is shutdown, and then restarted before it is fully cooled, the casings and rotor can be thermally distorted when the engine is restarted, causing rubs at some circumferential locations and not others.
Minimising the clearances has conventionally been by virtue of selecting the appropriate clearance for all engine operational points and transient variations. The nominal geometry is subject to manufacturing and build tolerances, thus to ensure a clearance around the whole rotor assembly accommodating the tolerances increases the overall tip clearance area.
Conventional manufacturing of the compressor components involves machining the whole stator casing and all the stator vane tips concentrically and axisymmetrically about the nominal engine or compressor centreline. The diameter of the machining operation is set to avoid any heavy rub that might damage the components.
Therefore, conventional manufacture of compressor (or turbine) components compromises the efficiency because each stage's machined diameter is axisymmetric about a single engine centreline, and the cold build clearance between rotor and stator of each stage is based on accommodating the worst case rub at any local point around the circumference.
U.S. Pat. No. 6,409,471 discloses a method of machining an inner surface of a shroud assembly extending generally circumferentially around a central axis of a gas turbine aircraft engine. The method includes determining pre-machined radial clearances during flight between tips of rotor blades in the engine and the inner surface of the shroud assembly at each of a plurality of circumferentially spaced locations around the shroud assembly. The inner surface of the shroud assembly is machined based on the pre-machined radial clearances to provide a generally uniform post-machined radial clearance during flight between the tips of the rotor blades and the inner surface of the shroud assembly at each of the circumferentially spaced locations around the shroud assembly.